The present invention relates to a method and system for attitude control of a spacecraft utilized when an inertial measurement unit is either permanently inoperable, temporarily unavailable, or otherwise not present.
Zero-momentum stabilized spacecraft typically include at least one inertial measurement unit (IMU) to measure angular rates of the spacecraft. The inertial measurement unit includes a plurality of gyroscopes to measure spacecraft angular rates about each of three axes. A processor internal to the IMU controls its operation and a power supply provides power to the IMU and its components.
Measurements from the IMU are utilized to maintain the spacecraft pointing in a desired direction to permit the spacecraft to carry out its mission and to permit control of the spacecraft. In the case of an earth orbiting communications satellite, the IMU is utilized to maintain one face of the satellite pointed toward the earth. Failure of the IMU can result in mispointing of a spacecraft and its antennas, disrupting the spacecraft mission.
Known zero momentum control systems utilize the IMU to continuously measure rate for attitude determination and control. The angular rate information is used to continuously propagate the three-axis inertial attitude. Attitude sensor(s), such as earth and sun sensors, are employed to measure the spacecraft attitude and correct for errors in propagated attitude and estimate the gyro biases.
An earth sensor assembly (ESA) is used to measure roll and pitch attitude, while a sun sensor assembly (SSA) measures yaw attitude. The estimated attitude and rate are compared to a commanded attitude and rate, to generate control error signals. Resulting error signals are input to a proportional-integral-derivative (PID) controller that computes corrective control torque demands.
In the PID controller processing, the attitude error is multiplied by a proportional gain, the rate (derivative) error is multiplied by a derivative gain, and the attitude error is integrated and multiplied by an integral gain. For less precise pointing, a proportional-derivative (PD) controller may be used, where the integral term is deleted. The control torque demands generated by the PID controller for each of the spacecraft yaw, roll, and pitch axes are input to control logic that distributes the demands to actuators such as reaction wheels or thrusters to generate control torques. The control torques applied by the actuators reduce the attitude and rate errors, to maintain alignment between the spacecraft body coordinate frame and an earth-pointing coordinate frame.
The availability of continuous three-axis angular rate information from an IMU facilitates a zero-momentum control approach in which high-bandwidth attitude control loops may be closed on each of the yaw, roll, and pitch axes. Such high-bandwidth control loops are characterized by open-loop crossover frequencies between about 0.08 and about 0.2 rad/sec, transient response settling times of approximately 2 to 3 minutes, and good rejection of environmental and thruster-firing disturbance torques.
Known gyroless control systems use an earth sensor to measure roll and pitch, and a stored angular momentum along the pitch axis for indirect yaw control. The pitch momentum is typically stored in a momentum wheel assembly that is operated at near constant speed, with small commanded speed variations for pitch control. Roll and yaw control may be provided using magnetic torquers or thrusters.
For known systems, the bandwidth of the roll control is typically low, about an order of magnitude lower than the bandwidth of the zero-momentum roll control loop. The bandwidth of the roll control is constrained by the frequency of the spacecraft nutation mode. The frequency of this dynamics mode, that couples the roll and yaw axes, is determined by the bias pitch momentum and the spacecraft roll and yaw moments of inertia, according to the formula
xcex=h/{square root over (IxIy)}
where h is the pitch momentum bias, Ix is the yaw momentum of inertia, and Iy is the roll moment of inertia. For a spacecraft using a prior-art gyroless system, h=500 in-lb-sec, Ix=Iy=100,000 in-lb-sec2, and the nutation frequency is 0.005 rad/sec. The disadvantage of the low-bandwidth roll control is that about 20 to 30 minutes is required for the transient response to settle. Also, reduced disturbance rejection due to the lower bandwidth results in increased pointing errors due to environmental disturbance and thruster-firing torques.
Furthermore, the structure of the prior-art gyroless system is highly dissimilar to zero-momentum systems. First, these prior-art systems do not use reaction wheels for control. This is because these systems must apply external torques to control yaw, and reaction wheels can only apply internal torques. External roll and yaw torques, applied using magnetic torquers or thrusters, are used to control yaw by correcting the pointing of the spacecraft angular momentum vector in response to the sensed roll attitude error. The control system keeps the spacecraft momentum vector, nominally aligned with the spacecraft pitch axis, aligned with the orbit normal.
Second, as described herein, the roll control is implemented by applying primarily yaw control torque in response to the sensed roll error, and simultaneously applying yaw and roll torques to damp the nutation mode. This is in contrast to the zero-momentum roll control, which applies roll control torques in response to the sensed roll error. Furthermore, zero-momentum systems do not include nutation damping controllers, since the nutation mode, which may or may not exist depending on the momentum stored in the reaction wheels, is essentially eliminated by the closure of high bandwidth roll and yaw control loops.
The present invention provides a method and system for maintaining spacecraft pointing performance in the event of IMU failure. The present invention addresses the limitations of known gyroless spacecraft attitude control systems. In particular, the present invention provides a method for gyroless control that uses reaction wheels and retains a high degree of similarity to standard zero-momentum systems. Specifically, the gyroless roll control may use a standard PD controller that provides the same loop bandwidth as is typical of zero-momentum systems. Furthermore, the gyroless system may use the same reaction wheel torque distribution logic as a standard zero-momentum system. The modifications needed to provide gyroless yaw control may include simple add-ons to a zero-momentum system, and may be implemented with minimal changes to the flight software that operates in the spacecraft on-board processor. Because of its close similarity to zero-momentum control systems, a gyroless system according to the invention is well suited for incorporation in a hybrid system that can either operate in a gyro-based or gyroless mode. Such a system is described in patent application TBS, where the gyroless mode is automatically invoked to maintain payload pointing in the event of a failure of the inertial measurement unit.
Along these lines, the present invention provides a method for maintaining three-axis control of a geosynchronous spacecraft without body angular rate measurements and using reaction wheel assemblies. The method utilizes earth sensor assembly angle measurements for high-bandwidth roll and pitch control. A positive pitch momentum bias is stored in the reaction wheel assemblies. A gyroscopic feedforward torque is applied to rotate the reaction wheel assembly momentum in a yaw/roll plane of the spacecraft at orbit rate. A dynamic mode is damped based on earth sensor assembly roll measurements. The dynamic mode couples the yaw and roll axes and results from the pitch bias momentum, applying the gyroscopic feedforward. torque, and the high-bandwidth roll control.
The present invention also provides a method for gyroless control of a spacecraft. A reaction wheel assembly momentum on each axis of the spacecraft and earth sensor assembly roll and pitch angles are input to a controller. Earth sensor assembly angles and roll and pitch reaction wheel assembly momentum are processed through high bandwidth controllers for roll and pitch control. Earth sensor assembly roll angle and reaction wheel assembly roll and yaw momentum are processed for yaw control. Reaction wheel assembly torque demands are determined and an output generated with the controller to effect the reaction wheel assembly torque demands. Reaction wheel assemblies are actuated to effect the torque demands.
Additionally, the present invention provides a system for gyroless control of a spacecraft. The system includes high-bandwidth control loops operable to utilize earth sensor assembly angle measurements for roll and pitch control. The system also includes reaction wheel assemblies operable to store a positive pitch momentum bias and to receive a gyroscopic feedforward torque to rotate reaction wheel assembly momentum in a yaw/roll plane of the spacecraft at orbit rate. Additionally, the system includes a controller operable, based on earth sensor assembly roll measurements, to damp a dynamic mode that couples yaw and roll axes and that results from applying pitch momentum bias, the gyroscopic feedforward torque, and the high-bandwidth roll control.
Furthermore, the present invention provides a system for gyroless control of a spacecraft attitude. The system includes a controller operable to receive a reaction wheel assembly momentum on each axis of the spacecraft and earth sensor assembly roll and pitch angles. The controller determines reaction wheel assembly torque demands and generates output to effect the reaction wheel assembly torque demands and processes earth sensor assembly roll angle and reaction wheel assembly roll and yaw momentum for yaw control. A plurality of high bandwidth controllers are operable to process earth sensor assembly angles and roll and pitch reaction wheel assembly momentum for roll and pitch control.
Still further, the present invention provides a spacecraft that includes a plurality of reaction wheel assemblies, an earth sensor assembly, and a controller operable to carry out any of the methods described herein.
Still other objects and advantages of the present invention will become readily apparent by those skilled in the art from a review of the following detailed description. The detailed description shows and describes preferred embodiments of the present invention, simply by way of illustration of the best mode contemplated of carrying out the invention. As will be realized, the present invention is capable of other and different embodiments and its several details are capable of modifications in various obvious respects, without departing from the present invention. Accordingly, the drawings and description are illustrative in nature and not restrictive.